A method and system for autonomous deployment of a satellite constellation

Abstract

The application discloses a satellite constellation autonomous deployment method and system, belongs to the field of space orbit control technology, and solves problems of long deployment period, high personnel investment, slow collision avoidance response, increased orbit safety risk and the like caused by a traditional satellite deployment method. The method comprises the following steps: constructing a constellation network according to satellite initial orbit parameters, different stage orbiting time and orbit parameters; outputting the best deployment strategy of each stage of the constellation through a genetic algorithm; realizing the adjustment of the ascending node right ascension and the phase to the target position together while realizing the deployment in place through coupling equations, and constructing a firing constraint; generating a variable orbit plan according to the thrust and the initial orbit of the satellite to be deployed; autonomously generating a variable orbit instruction according to the variable orbit plan; judging the variable orbit progress according to on-orbit telemetry, and adjusting the strategy in a closed loop; and realizing the collaborative deployment of satellites on multiple orbit planes, saving fuel and time cost. The application is particularly suitable for the collaborative deployment and autonomous management of large-scale satellite constellations on multiple orbit planes.

Description

Technical Field

[0001] This invention belongs to the field of aerospace orbit control technology, specifically relating to a method and system for autonomous and collaborative deployment of low-Earth orbit constellations. Background Technology

[0002] With the mass production and large-scale launch of satellites, the large number of low-Earth orbit (LEO) satellites leads to long deployment cycles. Increased constellation size also necessitates greater personnel investment. Traditional satellite orbital maneuvers rely on ground stations, requiring the generation of orbit change commands based on ground-based calculations of passing orbit data. This reliance on uploading commands from ground stations introduces time delays, resulting in slow collision avoidance responses and increased orbital safety risks. When both phase and right ascension of the ascending node need to be deployed simultaneously, the general strategy is to adjust the slower variable, the right ascension of the ascending node, first, followed by phase adjustment, which is not only time-consuming but also wasteful of fuel. Summary of the Invention To address the technical problems mentioned in the background section, this invention constructs an autonomous deployment system. Through a unified software system, it manages the deployment of multi-orbit constellations, enabling coordinated orbit changes across multiple orbital planes. Simultaneously, it generates orbit change plans and daily orbit change commands for the constellation, and updates the orbit change strategy based on telemetry feedback of daily progress, achieving coordinated constellation deployment and reducing personnel requirements. Furthermore, it enables simultaneous adjustment of the ascending node's right ascension and phase; and after orbit deployment, it achieves long-term satellite tracking, generating an on-orbit orbit maintenance strategy.

[0003] To achieve the above objectives, the present invention provides the following technical solution: In a first aspect, the present invention provides a method for autonomous deployment of a satellite constellation, the deployment method comprising the following steps: Step 1: Based on the determined constellation configuration and the initial orbital parameters of the satellites provided by the launch vehicle, construct the constellation network according to the different stages of orbit insertion time and orbital parameters; Step 2: Optimize the target performance using a genetic algorithm to output the best deployment strategy for each stage of the constellation; Step 3: Establish coupling equations. Through these coupling equations, while achieving high-altitude deployment, the right ascension of the ascending node and the phase are adjusted to the target position, and ignition constraints are constructed. Step 4: Based on the thrust and the initial orbit of the satellite to be deployed, generate an orbit change plan through ignition constraints; Step 5: Generate track change instructions autonomously based on the track change plan; Step 6: Based on on-orbit telemetry, determine the progress of the orbit change and make closed-loop adjustments to the strategy.

[0004] Furthermore, step 1 above specifically includes: Step 1.1: Define the number of orbital planes P and the total number of satellites in the constellation T, obtain the number of satellites in the orbital plane as S=T / P, and assign a unique code to each satellite in the constellation; Step 1.2: Input initialization parameters, including the number of orbital elements, phase interval, launch vehicle separation time, and initial orbital altitude after launch vehicle separation; The six elements of the orbit include the target orbital altitude, inclination, eccentricity, right ascension of the ascending node, argument of perigee, and mean perigee. Step 1.3: Based on the propulsion specific impulse and thrust, and the satellite's dry weight, calculate the orbital velocity increment through Hohmann transfer to complete the fuel model establishment, thereby realizing the constellation network construction.

[0005] Furthermore, the optimal deployment strategy in step 2 above includes the number of satellites to be deployed and the number of orbital poles after deployment.

[0006] Furthermore, the above coupled equations are constructed by taking the optimization of total adjustment time and total fuel consumption as the objective function, and phase, right ascension of the ascending node, and height coordination as constraints.

[0007] Furthermore, the ignition duration and number of ignitions per rail are determined through ignition constraints.

[0008] Furthermore, by controlling ignition constraints and the output ignition position, autonomous trajectory change planning is achieved, specifically as follows: During orbit ascent, the satellite eccentricity is first determined. When the eccentricity is small (e < 0.001), the latitude argument is selected for ignition. When the apogee ignition condition is met, the time when the satellite passes through the apogee is calculated as the propulsion ignition center point. The ascending node is adjusted to be perpendicular to the orbital plane, and the time at the ascending node is the ignition center point.

[0009] Furthermore, the specific steps for generating orbit change commands are as follows: extract the satellite number, orbital six-point number, and orbital UTC time from the telemetry data source code transmitted from the satellite, and output the ignition time for the entire day based on the ignition position.

[0010] Secondly, the satellite constellation autonomous deployment method described in this invention can be entirely implemented using computer software. Therefore, correspondingly, this invention also provides a satellite constellation autonomous deployment system.

[0011] Thirdly, the present invention also provides a computer-readable storage medium storing a computer program, which, when executed by a processor, performs the autonomous deployment method for a satellite constellation described in any of the preceding claims.

[0012] Fourthly, the present invention also provides a computer device, the device including a memory and a processor, wherein the memory stores a computer program, and when the processor runs the computer program stored in the memory, the processor executes a satellite constellation autonomous deployment method as described in any of the preceding claims.

[0013] The beneficial effects of this invention are as follows: This invention provides a method for autonomous deployment of satellite constellations. It manages the deployment of multi-orbit constellations through a unified software system, enabling coordinated orbit changes across multiple orbital planes. Simultaneously, it generates orbit change plans and daily orbit change commands for the constellations and updates the orbit change strategy based on telemetry feedback of daily progress, thus achieving coordinated constellation deployment and reducing personnel input.

[0014] Furthermore, this invention uses the optimization of total adjustment time and total fuel consumption as the objective function, and the coordinated adjustment of phase, right ascension of the ascending node, and altitude as constraints to construct a coupling equation, so that the right ascension of the ascending node and phase and orbital altitude can be deployed simultaneously through the coupling equation, saving constellation deployment time and fuel consumption.

[0015] Furthermore, this invention generates orbit control strategies based on real-time orbit changes, enabling autonomous selection of latitude argument and right ascension ignition at the ascending node, thereby achieving autonomous generation of orbit change plans.

[0016] Furthermore, this invention reduces personnel input and improves work efficiency by autonomously generating orbit control commands from the ground.

[0017] Furthermore, based on on-orbit telemetry, this invention autonomously judges the orbit control execution status and performs closed-loop strategy adjustments.

[0018] This invention is particularly suitable for the multi-orbital-plane collaborative deployment and autonomous management of large-scale satellite constellations. Attached Figure Description

[0019] To more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the specific embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.

[0020] Figure 1 This is a flowchart of a satellite constellation autonomous deployment method as described in an embodiment of the present invention; Figure 2 This refers to the constellation deployment schedule described in this embodiment of the invention; Figure 3 This describes the distribution of satellites on each orbital plane as described in this embodiment of the invention. Figure 4This is a view of satellite orbit transfer as described in an embodiment of the present invention. Detailed Implementation

[0021] The specific implementation details (such as experimental apparatus, operation procedures, data processing steps, and example parameters) of "a method for autonomous deployment of a satellite constellation" provided in this specification are primarily intended for illustrative purposes rather than limiting definitions, aiming to help those skilled in the art thoroughly understand the principles and implementation of the invention. However, those skilled in the art should understand that these details represent only one feasible embodiment, and the core concept of the invention can be fully realized through other technical means or alternative solutions not described in detail, without departing from its spirit and essence. Furthermore, the omission of details of conventional experimental methods and apparatus known in the art in the specification is to avoid redundant information interfering with the understanding of the innovation points. This does not mean that these known technologies are not required during implementation, and those skilled in the art should be able to supplement and apply them based on their professional knowledge.

[0022] The specific embodiments of the present invention will be further described in detail below with reference to the accompanying drawings. The following examples will help those skilled in the art to further understand the present invention, but do not limit the present invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention, and these all fall within the protection scope of the present invention.

[0023] Example 1, Combination Figure 1 This embodiment addresses the problems of long deployment cycles, slow collision avoidance response, increased orbital safety risks, and time-consuming and fuel-wasting practices inherent in existing orbital deployment strategies. It constructs an autonomous deployment system that manages multi-orbit constellation deployment through a unified software system. This system enables coordinated orbit changes across multiple orbital planes, generates orbit change plans and daily orbit change commands, and updates the orbit change strategy based on daily telemetry feedback, achieving coordinated constellation deployment and reducing personnel requirements. It also allows for simultaneous adjustment of the ascending node's right ascension and phase. Furthermore, it enables long-term satellite tracking after deployment, generating an on-orbit orbit maintenance strategy.

[0024] like Figure 1 As shown, the autonomous deployment method for satellite constellations includes the following steps: Step 1: Based on the determined constellation configuration and the initial orbital parameters of the satellites provided by the launch vehicle, construct the constellation network according to the different stages of orbit insertion time and orbital parameters; Step 2: Based on the constructed constellation network, optimize the target performance of the constellation deployment strategy using a genetic algorithm, and output the optimal deployment strategy for each stage of the constellation deployment. Step 3: Establish coupling equations. Through these coupling equations, while achieving high-altitude deployment, the right ascension of the ascending node and the phase are adjusted to the target position, and ignition constraints are constructed. Step 4: Generate an orbit change plan based on the thrust and the initial orbit of the satellite to be deployed; Step 5: Generate track change instructions autonomously based on the track change plan; Step 6: Based on on-orbit telemetry, determine the progress of the orbit change and make closed-loop adjustments to the strategy.

[0025] Example 2, Combination Figure 2 and Figure 3 This embodiment describes in detail each step of the autonomous deployment method for a satellite constellation described in Embodiment 1 above. Step 1: Based on the determined constellation configuration and the initial orbital parameters of the satellites provided by the launch vehicle, construct the constellation network according to the different stages of orbit insertion time and orbital parameters; Specifically: Step 1.1: Define the number of orbital planes P and the total number of constellation satellites T, then the number of satellites in the orbital plane can be obtained as follows: S =T / P, and assign a unique code to the satellites in the constellation, with the code ranging from 1 to T; Step 1.2: Input initialization parameters, including the number of orbital elements, phase interval, launch vehicle separation time, and initial orbital altitude after launch vehicle separation.

[0026] The six orbital roots are the core parameters used to define the shape and orientation of a satellite in space. They include the target orbital altitude (which determines the satellite's speed and period), inclination (the angle between the orbital plane and the Earth's equatorial plane, which determines the constellation's coverage), eccentricity (which describes the ellipticity of the orbit), right ascension of the ascending node (which determines the orientation of the orbital plane in space), argument of perigee (the angle from the ascending node to the perigee within the satellite's orbital plane), and mean perigee. These six parameters constitute the complete "six orbital roots," uniquely determining a satellite's orbit at a given moment.

[0027] The specific formula is as follows:

[0028]

[0029]

[0030] In the formula, The semi-major axis representing the target satellite orbit. Represents the target orbital altitude. Represents the Earth's radius; Represents the right ascension of the ascending node; Represents the near-point angle.

[0031] Phase spacing is used to define parameters of the constellation's internal structure, while the launch vehicle separation time and initial orbital altitude after separation are key parameters for defining the deployment sequence and initial state.

[0032] Step 1.3: Input the propulsion specific impulse and thrust, satellite dry weight, and use the Hohmann transfer to calculate the orbital velocity increment. This allows for the establishment of a fuel model and the completion of constellation network construction.

[0033]

[0034]

[0035]

[0036]

[0037] in, Represents the semi-major axis of the initial orbit. Represents the semi-major axis of the target orbit. Represents the orbital phasing speed increment. This represents the required fuel mass. Represents the initial mass of the spacecraft. For total increment, It is a comparison.

[0038] Step 2: Based on the constructed constellation network, optimize the target performance of the constellation deployment strategy using a genetic algorithm, and output the number of satellites to be deployed in each stage of the constellation, as well as the number of orbital elements after deployment; Specifically: First, define the decision space: in the j-th stage, deploy m satellites into orbit.

[0039] ,

[0040]

[0041] Multi-objective weighted sum

[0042] Selection, crossover, and mutation (swapping the deployment phases of two satellites, reallocating satellites to launches with more fuel) – that is, using genetic algorithms to optimize target performance to obtain the best deployment method, such as... Figure 2 and Figure 3 As shown.

[0043] In practical applications, the population is first initialized by randomly shuffling the constellation satellites and allocating launches according to payload capacity. A multi-objective weighted strategy is then used to assess fitness by minimizing fuel consumption and maximizing coverage. Next, a selection process is performed using a tournament selection method to preserve population diversity. Crossover is conducted at two points, randomly selecting two crossover points and swapping the launch segments of the two parent satellites between those points. After crossover, issues such as duplicate or missing satellites need to be addressed. Mutation involves swapping the positions of one satellite from each of the two launches.

[0044] Step 3: Establish the coupling equation. Through the coupling equation, while achieving the high-altitude deployment, the right ascension of the ascending node and the phase are adjusted to the target position. Specifically: Step 3.1: Based on the target orbit generated in Step 2 above, optimize the total adjustment time and total fuel consumption, and make coordinated adjustments in phase, right ascension of the ascending node, and altitude; (1) in, For rocket equations, For fuel consumption weighting, Total adjustment time weight, , To normalize the parameters and eliminate the influence of dimensions.

[0045] In practical applications, the target performance is optimized using a genetic algorithm, outputting the number of satellites to be deployed in each stage of the constellation, as well as the six orbital elements after deployment, which serve as the target orbits. The total adjustment time and total fuel consumption are used as the objective function, and the coordinated adjustment of phase, right ascension of the ascending node, and altitude are used as constraints to construct a coupling equation, as shown in formula (1), thereby obtaining the rate of change of right ascension of the ascending node and the rate of change of phase, and thus achieving coordinated adjustment of phase, right ascension of the ascending node, and altitude.

[0046] The rate of change of right ascension of the ascending node is:

[0047] in, For track inclination, The total velocity increment is adjusted for the right ascension of the ascending node.

[0048] The phase change rate is:

[0049] in, For the initial orbital period, This is a temporary orbital period.

[0050] Step 3.2: Input ignition constraints, which constrain the ignition duration per rail and the number of ignitions per rail.

[0051] Step 4: Generate an orbit change plan based on the thrust and the initial orbit of the satellite to be deployed; Specifically: Based on step 3 above, the ignition constraints and output ignition positions are determined to achieve autonomous orbit change planning. Specifically: First, the satellite eccentricity is determined. When the eccentricity is small (e < 0.001), latitudinal argument ignition is prioritized, i.e., the ignition center time is cyclically selected between 92° and 272° latitude arguments. Otherwise, apogee ignition is performed, and the satellite's apogee time is calculated as the propulsion ignition center point. After determining the ignition center time, an orbit change plan is generated based on the target orbit and current thrust.

[0052] Step 5: Generate track change instructions autonomously based on the track change plan; Specifically: Step 5.1: Extract the satellite number, orbital six-point number and orbital UTC time from the telemetry data source code transmitted from the satellite, and output the ignition time for the whole day according to the ignition position generated in Step 4.

[0053] Calculate the time of apogee:

[0054] in, It is the angle of approach, and is linearly related to time. It is the average angular velocity. It is the moment when the satellite passes through its closest point to Earth. For the current moment, For the near-point angle, Let be the orbital eccentricity.

[0055] At the apogee, the angle is closer to the perigee. Substituting into Kepler's equation:

[0056] Therefore, the angle of near point Time corresponding to apogee time:

[0057] The time of apogee is:

[0058]

[0059]

[0060] That is, the apogee time is the perigee time plus half an orbital period.

[0061] The formula for calculating the apogee time for the entire 24 hours of the day based on the time and angle of apogee corresponding to the initial orbit is as follows:

[0062]

[0063] Formula for calculating the argument of a fixed latitude:

[0064] The argument of latitude represents the angular position relative to the ascending node in the orbital plane, where, The perigee argument, True close angle.

[0065] To avoid ignition within the territory, choose In order to maintain a circular trajectory, symmetrical ignition is selected, i.e. .

[0066]

[0067]

[0068] )

[0069] Formula for the effect of small impulse on each number of roots:

[0070] Step 5.2: Determine the timing of two adjacent ignitions based on the orbital cycle. If the orbital cycle interval is met, generate orbital control commands and save the source code to a txt file, as well as output the orbital altitude change, fuel consumption, total ignition duration, mission start time, and mission end time.

[0071] Step 6: Based on on-orbit telemetry, determine the progress of the orbit change and make closed-loop adjustments to the strategy.

[0072] Specifically: By transmitting telemetry data from satellites, daily propulsion progress, as well as propulsion altitude and fuel consumption, a full-day orbit control report is generated, which in turn determines the progress of orbit changes and enables closed-loop strategy adjustments.

[0073] Example 3, Combination Figures 1 to 4 This embodiment provides a specific implementation method to comprehensively explain the autonomous deployment method for satellite constellations described in the above embodiments.

[0074] The process of autonomous deployment of satellite constellations is as follows: Figure 1 As shown, it includes the following steps: Step 1.1: With the number of orbital planes P=6 and the total number of constellation satellites T=60, the number of satellites in the orbital plane is S=10. Create 60 random individuals, numbered 1~60.

[0075] Step 1.2: Target orbital altitude 1100km, inclination 55°, eccentricity 0.001, satellite-launch separation time [2025-10-01, 2025-12-01, 2026-01-01, 2026-02-01, 2026-03-01], initial orbital altitude after satellite-launch separation [500, 550, 600, 550, 600]. Thrust specific impulse is 1200s, total satellite weight is 200kg.

[0076] After initializing the parameter settings, a fuel model is established using the Hohmann transfer.

[0077] Step 2.1: Use a genetic algorithm to configure parameters, with a population size of 100, 200 iterations, a crossover probability of 0.8, a mutation rate of 0.3, optimize coverage and fuel, and generate the deployment strategy shown in Table 1 (six roots of the constellation). Figure 2 What is shown is the constellation deployment progress. Figure 3 What is shown is the distribution of satellites in each orbital plane.

[0078] Table 1

[0079] Step 3.1: Optimize the total adjustment time and total fuel consumption as the objective function, and use phase, right ascension of the ascending node, and height as constraints; Step 3.2: Confirm the satellite's ignition capability, and input the ignition duration per orbit and the number of ignitions per orbit; Step 4.1: Solve the optimization problem and generate a trajectory change plan by outputting the ignition position for each step; Step 5.1: Extract the six data points and time from satellite telemetry, calculate the ignition time for the whole day, generate a report, and select one day for the ignition time, as shown in Table 2 (Daily Progress). Table 2

[0080] Step 6.1: Download telemetry data from the satellite, collect daily progress statistics as shown in Table 2, and draw a schematic diagram of orbital changes, as shown below. Figure 4 As shown.

[0081] Example 4: Any of the above-described satellite constellation autonomous deployment methods can be fully implemented using computer software. Therefore, this example provides a satellite constellation autonomous deployment system.

[0082] Example 5: This example provides a computer-readable storage medium storing a computer program, which, when executed by a processor, performs a satellite constellation autonomous deployment method as described in any of the above examples.

[0083] Those skilled in the art will understand that implementing all or part of the processes in the above embodiments can be accomplished by a computer program instructing related hardware. The program can be stored in a computer-readable storage medium, and when executed, it can include the processes of the embodiments of the above methods. The storage medium can be a magnetic disk, optical disk, read-only memory (ROM), random access memory (RAM), flash memory, hard disk drive (HDD), or solid-state drive (SSD), etc.; the storage medium can also include combinations of the above types of memory.

[0084] Example 6: This example provides a computer device, which includes a memory and a processor. The memory stores a computer program. When the processor runs the computer program stored in the memory, the processor executes a satellite constellation autonomous deployment method as described in any of the above examples.

[0085] This embodiment provides a computer device. This part of the hardware device is a general model and is not shown in the figure. The system includes a processor and a memory. The processor and the memory can be connected by a bus or other means. The memory, as a non-transitory computer-readable storage medium, can be used to store non-transitory software programs, non-transitory computer-executable programs and modules, as well as corresponding program instructions / modules. The processor executes various functional applications and data processing by running the non-transitory software programs, instructions and modules stored in the memory, so as to realize the autonomous deployment method and steps of a satellite constellation in the above method embodiment.

[0086] The memory may include a program storage area and a data storage area, wherein the program storage area may store the operating system and applications required for at least one function; the data storage area may store data created by the processor, etc. Furthermore, the memory may include high-speed random access memory and may also include non-transitory memory, such as at least one disk storage device, flash memory device, or other non-transitory solid-state storage device. In some embodiments, the memory may optionally include memory remotely located relative to the processor, which can be connected to the processor via a network. Examples of such networks include, but are not limited to, the Internet, corporate intranets, mobile communication networks, and combinations thereof.

[0087] One or more modules are stored in the memory. When the processor executes, it performs the method steps in the embodiments. In this way, the inventive purpose of the present invention can be achieved through the method, apparatus and process of the present invention. The specific details of the above-mentioned computer device can be understood by referring to the relevant descriptions and effects in the embodiments, and will not be repeated here.

[0088] The above description of the technical solution provided by the present invention through several specific embodiments is intended to highlight the advantages and benefits of the technical solution provided by the present invention. However, the above-described specific embodiments are not intended to limit the present invention. Any reasonable modifications and improvements to the present invention, reasonable combinations of implementation methods and equivalent substitutions based on the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A method for autonomous deployment of a satellite constellation, characterized in that, The method is as follows: Step 1: Based on the determined constellation configuration and the initial orbital parameters of the satellites provided by the launch vehicle, construct the constellation network according to the different stages of orbit insertion time and orbital parameters; Step 2: Optimize the target performance using a genetic algorithm to output the best deployment strategy for each stage of the constellation; Step 3: Establish coupling equations. Through these coupling equations, while achieving high-altitude deployment, the right ascension of the ascending node and the phase are adjusted to the target position, and ignition constraints are constructed. Step 4: Based on the thrust and the initial orbit of the satellite to be deployed, generate an orbit change plan through ignition constraints; By controlling ignition constraints and outputting ignition positions, autonomous trajectory change planning is achieved, specifically as follows: During orbit ascent, the satellite eccentricity is first determined. When the eccentricity is small (e < 0.001), latitude argument ignition is prioritized. When the apogee ignition condition is met, the time when the satellite passes through the apogee is calculated as the propulsion ignition center point. The ascending node is adjusted to be perpendicular to the orbital plane, and the time at the ascending node is the ignition center point. Step 5: Generate track change instructions autonomously based on the track change plan; Specifically: The satellite number, orbital six-point number, and orbital UTC time are extracted from the telemetry data source code transmitted from the satellite, and the ignition time for the whole day is output according to the ignition position. Step 6: Based on on-orbit telemetry, determine the progress of the orbit change and make closed-loop adjustments to the strategy.

2. The method for autonomous deployment of a satellite constellation according to claim 1, characterized in that, Step 1 is as follows: Step 1.1: Define the number of orbital planes P and the total number of satellites in the constellation T, obtain the number of satellites in the orbital plane as S=T / P, and assign a unique code to each satellite in the constellation; Step 1.2: Input initialization parameters, including the number of orbital elements, phase interval, launch vehicle separation time, and initial orbital altitude after launch vehicle separation; Step 1.3: Based on the propulsion specific impulse and thrust, and the satellite's dry weight, calculate the orbital velocity increment through Hohmann transfer to complete the fuel model establishment, thereby realizing the constellation network construction.

3. The method for autonomous deployment of a satellite constellation according to claim 1, characterized in that, The optimal deployment strategy includes the number of satellites to be deployed and the number of orbital poles after deployment.

4. The method for autonomous deployment of a satellite constellation according to claim 1, characterized in that, The total adjustment time and total fuel consumption are used as the objective function, and the phase, right ascension of the ascending node, and height coordination are used as constraints to construct a coupled equation.

5. The method for autonomous deployment of a satellite constellation according to claim 1, characterized in that, The ignition duration and number of ignitions per rail are determined by ignition constraints.

6. A satellite constellation autonomous deployment system, characterized in that, The system is implemented based on a satellite constellation autonomous deployment method according to any one of claims 1-5.

7. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores a computer program that, when executed by a processor, performs an autonomous deployment method for a satellite constellation as described in any one of claims 1-5.

8. A computer device, characterized in that, The device includes a memory and a processor. The memory stores a computer program. When the processor runs the computer program stored in the memory, the processor executes a satellite constellation autonomous deployment method according to any one of claims 1-5.

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